Enhanced adhesion thermal barrier coating

ABSTRACT

A gas turbine engine including a compressor section, a combustor section fluidly connected to the compressor section, a turbine section fluidly connected to the combustor section, and a plurality of gas path components exposed to a primary fluid flowpath through the compressor section, the combustor section and the turbine section. At least one of the gas path components includes an exterior facing surface, a lattice structure extending outward from the exterior facing surface, the lattice structure being integral to the exterior facing surface, and a thermal barrier coating adhered to at least a portion of the exterior facing surface and the lattice structure.

TECHNICAL FIELD

The present disclosure relates generally to thermal barrier coatings forgas powered turbine components, and more specifically to a system andprocess for enhancing adhesion between the thermal barrier coating andan underlying surface of a component for a gas powered turbine

BACKGROUND

In order to increase efficiencies of gas turbine engines, gas turbineengine manufacturers rely on extreme turbine inlet temperatures toprovide a boost to the overall engine performance. In some modern gasturbine engine applications, the gas path temperatures within theturbine exceed the melting point of the constituent materials from whichthe underlying components of the gas path are constructed. To addressthe extreme heat, cooling systems are used to cool the gas pathcomponents in the turbine.

One exemplary mechanism for cooling turbine gas path components is aceramic thermal barrier coating (TBC) adhered to an exterior surface ofthe gas path component. The presence of the thermal barrier coatingsignificantly reduces the operating temperature of the component andallows for lower cooling flow requirements.

SUMMARY OF THE INVENTION

In one exemplary embodiment a gas turbine engine component includes anexterior facing surface, a lattice structure extending outward from theexterior facing surface, the lattice structure being integral to theexterior facing surface, and a thermal barrier coating adhered to atleast a portion of the exterior facing surface and the latticestructure.

In another exemplary embodiment of the above described gas turbineengine component the lattice structure extends a first distance outwardfrom the exterior facing surface and wherein the first distance is lessthan a thickness of the thermal barrier coating.

In another exemplary embodiment of any of the above described gasturbine engine components the lattice structure is at least partially anartifact of a manufacturing process.

In another exemplary embodiment of any of the above described gasturbine engine components the manufacturing process is an additive metalmanufacturing process.

In another exemplary embodiment of any of the above described gasturbine engine components the exterior facing surface includes aplurality of through holes, and wherein the thermal barrier ischaracterized by not obstructing the through holes.

In another exemplary embodiment of any of the above described gasturbine engine components the exterior facing surface and the latticestructure have a combined surface roughness (Ra) in the range of100-600.

In another exemplary embodiment of any of the above described gasturbine engine components a surface roughness of an exterior surface ofthe thermal barrier coating is less than 100 Ra.

In another exemplary embodiment of any of the above described gasturbine engine components the combined surface roughness (Ra) is greaterthan 180.

In another exemplary embodiment of any of the above described gasturbine engine components the lattice structure is disposed across lessthan 100% of the exterior surface.

In another exemplary embodiment of any of the above described gasturbine engine components the component is one of a blade outer airseal, a blade and a vane.

An exemplary method for manufacturing a gas path component for a gaspowered turbine includes additively manufacturing a first componentstructure and a support structure, wherein the support structure isintegral to the first component structure, partially removing thesupport structure via a finishing process such that an artifact of thesupport structure remains integral to the first component structure, andapplying a thermal barrier coating to the first component structure,wherein the thermal barrier coating is adhered to the first componentstructure and the artifact of the support structure.

In another example of the above described method for manufacturing a gaspath component for a gas powered turbine applying the thermal barrierincludes applying a metallic bondcoat and applying a ceramic topcoat.

In another example of any of the above described exemplary methods formanufacturing a gas path component for a gas powered turbine additivelymanufacturing the first component structure and the support structureincludes constructing a lattice structure on at least one externalsurface of the first component structure.

In another example of any of the above described exemplary methods formanufacturing a gas path component for a gas powered turbine thethickness of the thermal barrier coating is larger than a height of theartifact of the support structure normal to the surface of the firstcomponent structure on which the artifact of the support structure isdisposed.

In another example of any of the above described exemplary methods formanufacturing a gas path component for a gas powered turbine partiallyremoving the support structure via the finishing process comprisesapplying the finishing process until a combined surface roughness of thefirst component structure and the artifact of the support structure hasa combined surface roughness (Ra) in the range of 100-600.

In another example of any of the above described exemplary methods formanufacturing a gas path component for a gas powered turbine partiallyartifact of the support structure is disposed across less than 100% ofthe exterior surface of the first component structure.

In one exemplary embodiment a gas turbine engine includes a compressorsection, a combustor section fluidly connected to the compressorsection, a turbine section fluidly connected to the combustor section,and a plurality of gas path components exposed to a primary fluidflowpath through the compressor section, the combustor section and theturbine section. At least one of the gas path components includes anexterior facing surface, a lattice structure extends outward from theexterior facing surface, the lattice structure being integral to theexterior facing surface, and a thermal barrier coating adhered to atleast a portion of the exterior facing surface and the latticestructure.

In another exemplary embodiment of the above described gas turbineengine the at least one of the gas path components includes one of ablade outer air seal, a blade, and a vane.

In another exemplary embodiment of any of the above described gasturbine engines the lattice structure is at least partially an artifactof an additive manufacturing process.

In another exemplary embodiment of any of the above described gasturbine engines a combined surface roughness of the exterior facingsurface and the lattice structure is a first magnitude and a combinedsurface roughness of the thermal barrier coating, the exterior facingsurface, and the lattice structure has a second magnitude, the secondmagnitude being less than the first magnitude.

These and other features of the present invention can be best understoodfrom the following specification and drawings, the following of which isa brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically illustrates an exemplary gas turbine engine.

FIG. 2 schematically illustrates an exemplary portion of a turbinegaspath.

FIG. 3 schematically illustrates the blade of FIG. 2 isolated from theturbine gaspath.

FIG. 4 schematically illustrates a cross sectional view of the blade ofFIG. 3.

FIG. 5 schematically illustrates an exemplary lattice structure on asurface of the blade of FIG. 2.

FIG. 6 schematically illustrates a partial cross sectional view of theexemplary lattice pattern of FIG. 5.

FIG. 7 schematically illustrates a build apparatus for constructing theblade of FIG. 2.

DETAILED DESCRIPTION OF AN EMBODIMENT

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five (5:1). Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (1066.8 meters). The flight condition of 0.8 Mach and35,000 ft (1066.8 m), with the engine at its best fuel consumption—alsoknown as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—isthe industry standard parameter of lbm of fuel being burned divided bylbf of thrust the engine produces at that minimum point. “Low fanpressure ratio” is the pressure ratio across the fan blade alone,without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressureratio as disclosed herein according to one non-limiting embodiment isless than about 1.45. “Low corrected fan tip speed” is the actual fantip speed in ft/sec divided by an industry standard temperaturecorrection of [(Tram ° R)/(518.7°R)]{circumflex over ( )}0.5. The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5 m/s).

Turbine engine components, alternately referred to as gas pathcomponents, that are exposed to the combustion products formed in thecombustor section 26 are exposed to extreme temperatures duringoperation of the gas powered turbine. In order to prevent damage to thegas path components, a ceramic thermal barrier coating is adhered tosurfaces of the gas path components that are exposed to the gas path.The ceramic thermal barrier coating reduces the operating temperature ofthe component and allows lower cooling flow requirements for activecooling systems used to cool the gas path components.

In some exemplary systems, ceramic thermal barrier coatings aresusceptible to breakage due to mechanical spallation. The breakage insuch examples is accelerated in regions with high surface curvatureand/or high heat load. By way of example, the leading edge of an airfoilis typically curved in such a manner.

With continued reference to FIG. 1, FIG. 2 schematically illustrates aturbine inlet 100 immediately aft of a combustor 110. Combustionproducts 112 are expelled from the combustor 110 along a gas path 102.Disposed within the gas path 102 are a vane 120 and a blade 130. Theblade 130 is supported by a disk 132 according to standard turbine bladeconstructions. Radially outward of the blade 130 is a blade outer airseal 140. Both the blade outer air seal 140 and the vane 120 aresupported by an engine case structure 150. Each of the surfaces of thevane 120, blade 130, and blade outer air seal 140 that are exposed tothe combustion products 112 in the gas path 102 are at least partiallycoated with a thermal barrier coating.

With continued reference to FIG. 2, FIG. 3 schematically illustrates theblade 130 of FIG. 2 isolated from the gas path structure. Similarly,FIG. 4 illustrates a cross sectional view of the blade 130 of FIG. 2along cross section lines A-A. The blade 130 includes a leading edge 210and a trailing edge 220. A root section 230 connects the blade 130 tothe disk 132 (illustrated in FIG. 2). A thermal barrier coating 240 isapplied to the leading edge 210 of the blade 130. In alternativeexamples, the thermal barrier coating 240 can be applied to the entireblade 130, instead of being limited to the leading edge 210. Someexample gas path components include holes disposed on the surfaceconnecting an internal cavity to the exterior of the gas path component.In such examples, the thermal barrier coating does not cover or impedethe through holes.

Due to the curvature of the leading edge 210, adherence of the thermalbarrier coating 240 to the exterior facing surface of the bladestructure is reduced under extreme heat loads, and mechanical spallationcan result. In order to increase the adherence of the thermal barriercoating 240 to the exterior facing surface of the blade structure, alattice structure (illustrated in FIG. 5) is disposed on the leadingedge 210 of the blade 130. The lattice structure protrudes outward fromthe exterior facing surface of the blade 130 and is covered by thethermal barrier coating 240.

With continued reference to FIGS. 2-4, FIG. 5 illustrates the latticestructure 400 positioned on the exterior facing surface of the blade130. FIG. 6 illustrates a cross sectional view of the lattice structure400 of FIG. 5 along view lines C-C. The lattice structure 400 is formedof a series of interconnected walls 410 extending outward from anexterior facing surface 420 of the blade 130. In the example latticestructure 400, each of the walls 410 extends approximately normal to theexterior facing surface 420 on which the lattice structure 400 isdisposed. In alternative examples, the lattice structure 400 can extendat an angle to the surface, with a component of the angle being normalto the exterior facing surface 420. While illustrated in FIG. 5 as aconnected pattern of hexagons, practical implementations can utilize anyinterconnected shape, and the lattice structure 400 is not limited tothe illustrated hexagonal shape.

A thermal barrier coating 430 is applied over the lattice structure 400.The thermal barrier coating 430 in this example is applied by firstapplying a metallic bondcoat and then applying a ceramic topcoat. Inalternative examples, alternative thermal barrier coating applicationtechniques can be utilized to similar effect. A depth 432 of the thermalbarrier coating 430 normal to the exterior facing surface 420 is lessthan a height 412 of the normal component of the lattice structure 400.The surface area generated by the sides of the lattice structure 400provides a greater surface for the thermal barrier coating to adhere toand reduces the occurrence of spallation.

In some examples, prior to application of the thermal barrier coating,the lattice structure 400 provides a surface roughness (Ra) of theexterior facing surface 420 in the range of 100-600 Ra. In otherexamples, the surface roughness provided by the lattice structure 400 isin the range of 180-600 Ra. As the thermal barrier coating 430 has adepth 432 greater than the height 412 of the lattice structure 400, theresultant surface roughness of the component after application of thethermal barrier coating 430 is significantly lower than the surfaceroughness of the exterior facing surface 420. The specific surfaceroughness of the thermal barrier coating 430 can be designed to a neededroughness according to known thermal barrier coating techniques. In someexamples, the resultant surface roughness of the thermal barrier coatingis less than 100 Ra.

With continued reference to FIGS. 2-6, FIG. 7 illustrates a completedblade 610 constructed in an additive metal manufacturing device 600.

Additive metal manufacturing typically utilizes lasers or electron beamsto sinter particles in a 2D powder bed. Parts, such as the blade 610,are made by successively sintering layers upwards to form the component.In order to create overhanging structures using the additive metalmanufacturing process, support material and frameworks must be utilizedto maintain the orientation of the part. When removed from the part, thesupport structures leave remnant structures that are attached to, andintegral to, the constructed part. The remnant structures arealternately referred to as artifacts. In existing systems the artifactsare removed in a finishing process.

In the illustrated example, the blade 610 is constructed with a leadingedge 620 approximately parallel to a base 602 of the additive metalmanufacturing device 600. Support structures 630 extend upward andcontact/support the blade 610 as it is being constructed. As will beunderstood by one of skill in the art, the support structures 630 areconstructed as a part of the additive metal manufacturing process.

In the illustrated example, the blade 610 and support structures 630 areconstructed layer by layer beginning at the base and extending in adirection indicated by the arrow 640. Each of the support structures 630contacts only the leading edge region of the blade 610. After the blade610 is constructed via the additive metal manufacturing process, thesupport structures 630 are removed from the blade 610 via a finishingprocess. By halting the finishing process while a portion of the supportstructure 630 remains connected to the leading edge, the artifacts fromthe support structure 630 form the lattice structure described above.One of skill in the art, having the benefit of this disclosure, willunderstand methods by which to finish the blade 610 to a desired surfaceroughness such that an appropriately sized lattice structure remains.

In alternative examples, such as those where a lattice structure isdesired across the entire surface of the blade 610, the supportstructures 630 can extend across the entire blade surface. In yetfurther alternative structures, a lattice structure can be constructedalong any desired exterior surface of the blade 610 during the additivemetal manufacturing process independent of the support structure.

While described above with general reference to application of a thermalbarrier coating to a leading edge of a turbine blade, one of skill inthe art will understand that the above described lattice structureapplied to an exterior facing surface of a component can be utilized toincrease adhesion of a thermal barrier coating to the exterior facingsurface of any component and any surface, and is not limited to aleading edge of a turbine blade.

It is further understood that any of the above described concepts can beused alone or in combination with any or all of the other abovedescribed concepts. Although an embodiment of this invention has beendisclosed, a worker of ordinary skill in this art would recognize thatcertain modifications would come within the scope of this invention. Forthat reason, the following claims should be studied to determine thetrue scope and content of this invention.

The invention claimed is:
 1. A method for manufacturing a gas pathcomponent for a gas powered turbine comprising: additively manufacturinga first component structure and a support structure, wherein the supportstructure is integral to the first component structure; partiallyremoving the support structure via a finishing process such that anartifact of said support structure remains integral to said firstcomponent structure by applying the finishing process until a combinedsurface roughness of the first component structure and the artifact ofthe support structure has a combined surface roughness (Ra) in the rangeof 100-600; and applying a thermal barrier coating to said firstcomponent structure, wherein said thermal barrier coating is adhered tosaid first component structure and said artifact of said supportstructure.
 2. The method of claim 1, wherein applying said thermalbarrier comprises applying a metallic bondcoat and applying a ceramictopcoat.
 3. The method of claim 1, wherein additively manufacturing thefirst component structure and the support structure comprisesconstructing a lattice structure on at least one external surface of thefirst component structure.
 4. The method of claim 1, wherein thethickness of the thermal barrier coating is larger than a height of theartifact of the support structure normal to the surface of the firstcomponent structure on which the artifact of the support structure isdisposed.
 5. The method of claim 1, wherein the artifact of said supportstructure is disposed across less than 100% of the exterior surface ofthe first component structure.